G0753646.pdf

IOSR Journal of Mechanical and Civil Engineering (IOSR-JMCE)

e-ISSN: 2278-1684,p-ISSN: 2320-334X, Volume 7, Issue 5 (Jul. - Aug. 2013), PP 36-46 www.iosrjournals.org

www.iosrjournals.org 36 | Page

Technical Development of Design & Fabrication of an Unmanned

Aerial Vehicle

Md. Fazlay Rabbey1, Easir Arafat Papon

2, Anik Mahmood Rumi

3,

Hafez Md. Monerujjaman4, Farhan Hasan Nuri

5

1,2,3,4,5(Aeronautical Engineering Department, Military Institute of Science & Technology, Bangladesh)

Abstract: UAV (Unmanned Aerial Vehicle) is an air vehicle which is largely used for surveillance, monitoring,

reconnaissance, data relay, and data collection or to enter the area which is not safe for human i.e. flood

affected or virus affected area. This paper represents the unique design of such an UAV which designed at

MILITARY INSTITUTE OF SCIENCE & TECHNOLOGY to participate in an international competition SAE Aero Design West-2013. As per competition requirement empty weight of the UAV must be less than 2 lb and

must fly with payload as heavy as possible for good scoring. Initially, the model of the UAV was tested in wind

tunnel and the test data showed that the model aircraft performance was capable enough for flying and covering

an area specified in the competition. Subsequently, an actual aircraft was fabricated of that model and flight

tested which proved the match with theoretical, statistical and experimental data that was obtained from wind

tunnel test, wing tip test, tensile test of manufacturing material and CFD (Computational Fluid Dynamics) flow

simulation over the aerofoil.

Keywords: Unmanned aerial vehicle (UAV), NACA 4412, unique design UAV.

I. INTRODUCTION Birds can fly in sky due to their inherent characteristics. They are the dominant in the sky. It is human

characteristics to dominate or rule over other. So they tried their best to fly in sky from dawn of civilization.

And they got it through the success of Wright brother’s in 1903. Subsequently many modernization and

invention had done through 19th century and now a day’s fly without pilot is one of the most important topics of

study worldwide for national, international, military purposes under the banner of Unmanned Aerial Vehicle

(UAV). Present paper discusses the systematic design, data analysis, different property calculations and then

manufacturing of such type of UAV with low cost which successfully flew in the sky in USA [as a part of

NASA-SAE competition at California 12-14 April 2013]. The project team secured 19th position in this contest.

II. GOAL OF THE PROJECT The goal of the project was to develop and exhibit a practical method of building a unique design UAV

and also take off, cruise and land it within the specified area given in the competition. This goal was being met

by engineering basics such as aerodynamics, propulsion, structural analysis, control system etc. The UAV was

launched by hand without the use of a runway for takeoff. Detail design was conducted using SolidWorks

software and then different parts of the prototype were manufactured from XPS foam, depron foam and other

materials using heat cutter machine. Ultimately, an airworthy aircraft was to be fabricated to meet competition

requirements.

III. CONCEPTUAL DESIGN The objective of conceptual design is to design an UAV based on the analysis of mission goals,

requirements and the design constraints set out in the competition. This was done through step by step analysis

of seven pivot points suggested by Raymer[7]. The expected result from this was to obtain a high performance

design that maximizes the overall flight score in the competition:

Final Flight Score = i x (Best Flight Score) (1)

Where, i = 1+ (A0-40%) x .25 (2)

Successful Flight

& A0 = (3)

Successful Flight + Missed Flight

Again, Flight Score= (2 – EW) x PF x 120 (4)

Where EW = Empty Weight in pounds, and

Technical Development Of Design & Fabrication Of An Unmanned Aerial Vehicle


Payload Weight

& PF = (5)

Payload Weight + Empty Weight

Using the formula it was ensured that for good performance empty weight was to be as small as possible whereas

carrying payload should be large. So, the aircraft configurations were selected in such a way that the empty

weight was small and it could fly with heavy payload. The aircraft configurations were selected based on the

following criteria [6]:

a) The conventional configuration was chosen over the flying wing, biplane etc. due to its easy of

manufacturing, less interference drag than biplane, high maneuverability, lightweight, flexibility of design and good L/D with much more reliability.

b) High wing aircraft has advantages over low and mid wing aircraft. Because high wing aircraft has the

increased lifting surface area at the centre of the wing where in other designs the wing attached to the fuselage

this lifting surface is lost. In high wing less Interference drag exists between the wing and fuselage.

c) Single tractor configuration was selected over the single pusher, double tractor and push-pull configuration

because it faces undisturbed flow while the pusher faces disturbed and vortex flow over the body and wing

respectively. For the single tractor also results in the overall lighter aircraft.

d) Conventional tail is selected over V-Tail, T-Tail, Tailless delta, Cruciform tail because it is easy to install and

provides robust and sufficient stability to the aircraft while in T-Tail weight penalty increased and in V-Tail

control authority of pitch and yaw is reduced.

The concept reflects the following preliminary design [6]:

Fig. 1: preliminary design by hand sketching

IV. DETAIL DESIGN OVERVIEW

Before doing main design the thing kept in mind was the given restrictions in the competition. Every

year SAE hosts a competition to design, construct and fly 3 types of aircraft i.e. micro class, advance class and

regular class. The team participated in micro class aircraft competition. Restrictions in this class were:

a) No lighter-than-air or rotary wing aircraft.

b) Maximum dimension of any parts can be 24 inch because all components for operation of the aircraft must be

put into a padded foam carrying case and inside dimension of Carrying Case should not exceed 24 inches by 18

inches by 8 inches (24"×18"×8").

c) The use of lead in any portion of the aircraft (payload included) is strictly prohibited.

d) Micro class aircraft can use electric motor only.

e) Aircraft must be capable of carrying and fully enclosing a rectangular block measuring 2 inches by 2 inches

by 5 inches (minimum size). The payload must consist of a support assembly and payload plates.

f) The aircraft must be capable of flight worthy without carrying a payload.

g) The aircraft must be able to take-off, complete a circuit and then safe land.

h) The UAV must be an empty weight of less than 2 lbs.

i) Maximum three (3) minutes to accomplish a successful takeoff.

j) Maximum three (3) minutes to assembly the aircraft from the padded foam carrying case in three (3) minutes

by two (2) people.

By keeping in mind restrictions the detail design [6] was conducted in SolidWorks and its basic layout is shown

in the following:



Fig. 2: 3D model of the UAV

Fig. 3: joint mechanism and attaching area of different parts

Fig. 4: all internal and external components of the UAV



V. DETAIL DIMENSIONS The designed UAV was performed a better aerodynamic efficiency when the flow passes through over

it during flight. The detail dimension of the UAV is as follows:

Table 1: Detail dimensions DETAIL DIMENSION HORIZONTAL STABILIZER AILERON

Airfoil Flat plate Span 10 in

WING Span 15 in % of chord 25%

Span 40 in Chord 5 in Max. δa 17.5°

Chord 6.3 in Area 75 in2 RUDDER

Aspect ratio 6.35 Incidence 2° Span 7 in

Taper ratio 1 Thickness 0.2 in Thickness 0.2 in

Wing area 252 in2 VERTICAL STABILIZER % of tip chord 45.5%

Airfoil NACA 4412 Airfoil Flat plate Max. δr 17.5°

Static margin (without payload)

26.5% Root Chord 5 in ELEVATOR

Static margin (with payload)

12.2% Tip chord 3.45 in Span 15 in

Incidence 3° Span 8.31 in % of chord 40 %

FUSELAGE Taper ratio 0.69 Max. δe (down) 17.5°

Length 23.22 in Thickness 0.2 in Max. δe (up) -17.5°

Max Width 3.2 in Mean Chord 4.27 in Thickness 0.2 in

VI. CALCULATIONS This step includes3D drag analyses, performance analysis, weight build up analysis & stability margin

calculation.

6.1 DRAG ANALYSIS INCLUDING 3-D DRAG EFFECT

The drag effects was analyzed through commercially available software like CFD, JAVAFOIL. By

analyzing different test results from these softwares, wind tunnel and historical data [11][1][4] the final finding was NACA 4412 aerofoil to meet the desired objectives. The aircraft drag polar was determined by following

[5]: CD = CD(min)+ K’CL2+ K’’(CL-CL(min))

2 (6)

Table 2: Ambient flight condition Ambient flight condition

Flight speed (Take-off) 10.72 ms-1

Air density at sea level 1.225kgm-3

Air pressure at sea level 1 atm

Sea level temperature 15°C

Kinematic viscosity of air, v 1.4607×10-5

m2/s

Gas constant, R 287JKg-1K

-1

Specific heat ratio, γ 1.4

For laminar flow: Cf,(lam)=1.328/(Re)0.5 (7)

When the BL is tending to turbulent: Cf,(tur)=0.455/((log10Re)2.58+(1+0.144M2)0.65) (8)

Table 3: Area & length of different parts Parts Planform area, SRef(in

2) Wetted area, SWet(in

2) Reference length (in)

Fuselage 60.47 274.67 23.22

Horizontal Tail 75 150 5(MAC)

Wing 252 504 6.3 (MAC)

Vertical Tail 0 67.3 3.17(MAC)

Fuselage: Re=VL/v=432806, assume BL is turbulent. Mach number: M=V/ (γRT)0.5 =0.0315

Co-efficient of friction: Cf,(turb)={0.455/(log10Re)2.58}-(1700/Re)=1.33×10-3

For the aircraft: FR = Fuselage fineness ratio = Fuselage length/width = 7.26..

Fuselage form factor: FFF = 1 + 60/(FR)3 + 0.0025 FR =1.1749

So, for fuselage: CD min = FFFCfSWet/SRef = 0.0071

Wing: For wing, Re = VC/v=117413.

Wing form factor: FFW = [1 + L(t/c) + 100(t/c)4] R ,Where L is the airfoil thickness location parameter (L = 1.2

for the max t/c located at 0.3c and L = 2.0 for the max t/c < 0.3c) and R is the lifting surface correlation parameter. Thus L = 1.2. For a low speed, unswept wing R is approximately 1.05.



Since a wing Re =117413 could be treated as laminar.

As the BL is laminar, the wing: Cf = 1.328/(Re)0.5=0.00388.

So wing form factor: FFW = [1 + L(t/c) + 100(t/c)4] R =1.223 and minimum drag coefficient: CD(min) = FFWCfSWet/SRef =0.0095

Horizontal Tail: For horizontal tail: Re=VCHT/v=93185.Therefore we’ll assume the BL is laminar.

Cf = 1.328/(Re)0.5=4.35×10-3.

Form factor: FFHT = [1 + L(t/c) + 100(t/c)4] R =1.1

and minimum drag coefficient: CD(min) = FFWCfSWet/SRef =0.0096

Vertical Tail: For vertical tail: Re = VCVT(MAC)/v =59042. Therefore assume the BL is laminar.

FFHT = [1 + L(t/c) + 100(t/c)4] R =1.127

Cf = 1.328/( Re )0.5=5.47×10-3

CD(min) = FFW CfSWet/ Sref= 0.0123

Total CD(min) : The total CD(min) is the sum of all the components. Thus the total model CD(min) = 0.0385 based

upon a wing reference area of 252in2

Total Drag Expression: Assuming a span effectiveness factor e = 0.95 gives an induced drag factor K’ = 1/( AR e) = 0.0528. Assuming the viscous drag factor K’’ = 0.0211 is 40% of the induced drag factor [10]. The

section drag polar data was determined to find out using JAVAFOIL. The determined CL(min)=0.4. So, the total

drag expression is:

CD = 0.0385 + 0.0528CL2 + 0.0211(CL – 0.4)2 (9)

The model drag polar is shown on the following Fig. 5:

Fig. 5: drag polar

6.2 WEIGHT ESTIMATION

As a first approximation from historical data, empty weight fractions, We/W0=0.56 [5]

Again, Payload weight, WP= 1.1 lb. But, gross weight: W0 = WP+ We (10)

W0 {(1- (We/W0)} = WP

1.1

W0 = = 2.5lb = 11.12N

1- 0.56

6.3 VSTALL & CLT/O ESTIMATION

For the UAV high A.R. wing was chosen with wing span, b= 40in; chord, c= 6.3 in

So, A.R. = 6.35;Wing Area, S= 252 in2=0.162m2 Again, the used was Aerofoil NACA 4412 for which the

CLmax= 1.22 (from wind tunnel testing)

But, wing loading [5]: W/S= 0.5×air density×Vstall2×CLmax (11)

So, stall velocity, Vstall = 9.5 m/s = 31.13 ft/s

Again, take-off lift-coefficient: CLT/O = CLmax/1.21 = 1.22/1.21 = 1.01 (12)

6.4 ESTIMATING PERFORMANCE

The UAV weight with payload, W = 11.16 N (From Weight build-Up Section).

Again, CLT/o= 1.01. So, take-off velocity: VTO = [2 W/(S CLT/o)]½ =10.72m/s (13)

Table 4: Propulsion summary KV parameter of the motor 1200 Kv

Battery voltage 11.1V

Propeller Diameter × Pitch 7×6 in

Power 250 W



Propeller RPM: RPM=KV×Volt=13320rev/min=1395rad/s (14)

Assuming,ηprop=0.8. But, pitch velocity: Vpitch=RPM×pitch×ηprop=170m/s (15)

Thrust:T=Power/Vpitch=250/170=1.47N (16) Thrust/Weight (with payload) Ratio= 0.132

6.5 TAKEOFF DISTANCE ESTIMATION

We Know, Roll distance after launching [5]:

1.21 (W/S)

Sg= = 44m=144ft (17)

g×air density×(CL)max(T/W)

But, turn radius [5]:

6.96(Vstall)2

R= = 64m =210ft (18)

g

Fig. 6: during airborne [5]

Again, from the above Fig. 6: R (1-cos uOB) = hOB (19)

& if height of obstacle, hOB = 35 ft

So, uOB= 33.560

So, Airborne Distance: Sa= R sin uOB =210sin33.560 =116 ft (20)

Total Take-off Distance: S = Sa + Sg = (144+116)ft = 260 ft (21)

6.6 STABILITY ANALYSIS

Stability margin equation: S.M= (xnp- xcg)/c. A small S.M. in the range of +5 to +15% is desired make

the UAV more maneuvering or responsive to elevator deflection. For calculating stability margin first need to

calculate neutral point which was 3.26 in from the leading edge of the UAV. Then the aircraft static margin was

calculated 26.5% (without payload) and 12.2% (with payload) calculated. The center of gravity (CG) needs to

be placed close to the aerodynamic center (AC) and ahead of the vehicle neutral point for both empty weight

and with payload for stability. The aircraft C.G. is 25.3% of chord from leading edge of wing for empty weight and is 39.5% of chord from leading edge of wing with payload along axis of aircraft on the plane of symmetry.

The most forward point [2] of C.G of our aircraft is calculated 0.966 inch behind the leading edge after

calculating the aerodynamic centre position of wing which is determined from the concept that change of

moment w.r.t. AOA about A.C. is zero[3][5].

Table 5: Stability Parameters Clα,w CLα,H xac,w(in)

(From

leading

edge of

wing

xac,H(in)

(From

leading

edge of

wing

ηH SH/S dαH/dα

xnp(in)

(From

leading

edge of

wing)

xmf(in)

(From

leading

edge of

wing)

xcg(in)

(From leading edge of

wing)

(Without payload)

5.13 1.38 1.68 1.67 0.8 0.337 0.6 3.26 0.966 1.591

6.7 WEIGHT BUILD-UP ANALYSIS

Basically To keep the aircraft stable the center of gravity (CG) needs to be placed properly with respect to the neutral point. In the designed UAV CG was placed close to the aerodynamic center (AC) for both the

empty weight and with payload. For most of the aircraft CG should be located between 25-30% of the chord

length behind the leading edge. The calculated CG in the following table and it is realized that the CG is very

close to the desired location.



Table 6: Weight build-up

VII. TESTING As mentioned earlier, different types of software test as well as physical hardware test was done before

fabricating the UAV to find out suitable design which could meet desired objectives. The testing results are:

7.1 WTT

Wind tunnel testing was conducted in AF 100 Subsonic Wind Tunnel in the aerodynamics lab in MIST

to determine the aerodynamic efficiency of wing to compare the experimental value with that theoretical

analysis of historical data. The CL(max) for NACA 4412 in the testing approximately equals the result of in JavaFoil and CFD. From the test it was also assured that NACA 4412 airfoil used perform a good aerodynamic

efficiency during flight for the UAV.

Fig. 7: WTT of NACA 4412

7.2 CFD FLOW SIMULATION RESULTS

(a) (b) (c)

Fig. 8: plot of velocity vectors, CP Vs position graph & pressure contours in FLUENT

As can be seen, the velocity of the upper surface is faster than the velocity on the lower surface. On the leading

edge, a stagnation point is seen where the velocity of the flow is nearly zero. The fluid accelerates on the upper

surface as can be seen from the change in colors of the vectors. On the trailing edge, the flow on the upper

Total mass(lb)

1.404(without payload) Weight(lb) Position in 30% chord

Mass Moment about 30% chord Actual

CG(in) 1.591in(without payload from wing leading edge) 2.49in(with payload from wing leading edge)

Electric components

Motor 0.196 8.0 1.568

ESC 0.066 3.7 0.2442

Battery 0.256 4.1 1.0496

Receiver 0.04 -0.5 -0.02

Propeller 0.026 9.5 0.247

Wing Wing, aileron servos, welds & hardware

0.38 -0.61 -0.2318

Fuselage Fuselage, elevator & rudder

servos, welds & hardware

0.374(without payload)

1.477(with payload)

-3.5

-1.3

-1.309

-1.9201

Empennage Vertical tail, welds & hardware 0.022 -16.5 -0.363

Horizontal tail, welds & hardware 0.044 -17.4 -0.7656

Total mass(lb) 1.404(without payload) 2.51(with payload)

0.4194(without payload) -1.5(with

payload)



surface decelerates and converges with the flow on the lower surface. The lower curve of Fig. 8(c) is for the

upper surface of the airfoil and has a negative pressure coefficient as the pressure is lower than the reference

pressure. It is seen that there is a region of high pressure at the leading edge (stagnation point) and region of low pressure on the upper surface of airfoil. This is what we expected from analysis of velocity vector plot.

Table 7: Drag co-efficient calculation of NACA-4412 by CFD Zone name Pressure

force(N)

Viscous

force (N)

Total force(N) Pressure

co-efficient

Viscous

co-efficient

Total

co-efficient

Airfoil 0.2526493 0 0.2526493 0.0035888955 0 0.0035888955

Net 0.2526493 0 0.2526493 0.0035888955 0 0.0035888955

Table 8: Lift co-efficient calculation of NACA-4412 by CFD Zone name Pressure

force(N)

Viscous force

(N)

Total force(N) Pressure

co-efficient

Viscous

co-efficient

Total

co-efficient

Airfoil 86.163106 0 86.163106 1.2239511 0 1.2239511

Net 86.163106 0 86.163106 1.2239511 0 1.2239511

Fig. 9: Cl value determination in CFD

7.3 STRUCTURAL TEST

Testing was conducted to verify that XPS foam has the strength to perform the desired property.

Tension testing performed on the tension test machines at SM Lab in MIST and it was confirmed from the tested

technical data that the structure will not fail in air. From stress strain plot of the XPS it was verified that the

elastic limit can sustain up to 0.02 strain limit with corresponding stresses of 35 psi.

During take-off lift must be greater than weight. On the basis of this principle wing load testing is

obligatory. The wing tip is set at two ends & weighed down with some load at the centre of wing. The wing was totally broken down after placing 3.3 lb on the wing. So, the maximum bending moment the UAV can sustain is

5.5 lb-ft.

1.667ft

Fig. 10: elliptical loading over the wing semi span

The maximum stress in such a beam is at the wing root. Since elliptical loading was assumed, so we first

computed the lift per unit span at the root of the wing. This was done by integrating the lift distribution and

setting the value equal to half the desired weight of the plane with payload. This gave a lift per unit span at the

wing root of 13.77 N/m from the following eqn: 𝐿√{1− (2𝑦/𝑏)^2}𝑏/2

0𝑑𝑦 =𝑊/2 (22)

7.4 FLIGHT TEST

Many flight tests conducted through a long time from January to April 2013 and then achieved the

highest performance design to successfully participate in the competition. Followings were the flight testing

results:



Table 9: Flight testing Flight Testing

Schedule

Purpose Consequences

03-01-13 To check the ability to fly Unairworthy

16-01-13 To check the hand launching take-off Conducted takeoff safely

29-01-13 To check the responding of the control system Rapid response

01-02-13 To test the stability performance Unstable

07-02-13 To test the capability of safe landing Crashed before landing

13-02-13 To check the control surface deflection Assured

26-02-13 Get a hold of safe landing Safe landing

08-03-13 To check the maneuverability during 1800 turn Successfully maneuvered

15-03-13 Verify the behavior of aircraft during flight with payload as per

competition missions requirements

Had enough static margin as

mission desired

27-03-13 To scrutinize the aircraft stability(whether the c.g. position

changed or unaffected)

c.g. position not too much change

with varying payload

VIII. AVIONICS The avionics control system of UAV relates to mainly controlling the craft on air which mainly

includes navigation, IMU sensor, GPS units, flight management, altimeters, servos etc. The servos used to

design the UAV are having: dimension 0.86 x 0.4 x 0.79 in; operating voltage 4.8V; stall Torque 0.6 kg/cm and

speed 0.1 sec/60 degree.

Single battery system with battery eliminator circuits used with a red shunt plug was positioned

between the ESC and motor. Lithium Polymer battery was used as the power supply because they worked well

and hold a large amount of current which was lighter than nickel metal and Ni-Cad batteries. The used Li-Po

battery can supply 11.1V and a current rating of 1350mAh. 3 cells Li-Po was used which weighted 115 gm in

total. Using 3 cells was advantageous as it was not to allow the battery voltage to drop too low. Using 3 cells

lithium polymer it was recommended to land the UAV when the voltage rating drops to 9V. Features of the

battery were: Cell Voltage: 3.6-3.7V (nominal), 3cells making 11.1V; Capacity: 1350mAh; Energy by Weight :

120Wh/L; Energy by Volume: 250Wh/L; Discharge characteristics: 35C; Temperature range: -200C to +600C;

Preferred Charge Methods: Constant Current/ Constant Voltage

The ESC used in the UAV operated at a Current Rating of 3A. The communication and navigation

system consisted of a transmitter and a receiver which operated based on the transmission and reception of radio

frequencies [8]. By this system the control surface movement and RPM was controlled from the ground. A

receiver was kept in UAV which received RF signals, modulate the signal and sent command to the control surfaces and motor according to instruction given from the ground. For communication system the used

transmitter was FS-TH9X 2.4GHz 9CH which operated at 2.4GHz frequency. As the transmitter module had 9

channels so 9 different commands could be sent at a time. Features of transmitter module: Channels: 9 channels

TH9X; Model type: Helicopter, Airplane and Glider; RF Power: Less than 20db; Modulation: GFSK; Code

type: PPM/PCM; Sensitivity: 1024; LCD type: 128*64 dot; Low voltage warning: Yes; DSC port: Yes; Charge

port: Yes; Power: 12VDC; Weight: 680gm.

8.1 CONTROL SYSTEM

The pilot controls the aircraft by remote during flight according to the following block diagram:



8.2 OPERATION CYCLE

During the competition the operation cycle was given through which the aircraft must fly with safe

landing for scoring. There was two option of operation cycle. The team took option 1 which is shown in the

following Fig. 11. Once takeoff occurs, it was necessary for aircraft in airborne and fly past a first cone before

turning approximately 180 degrees in heading, fly past a second cone & turning another 180 degrees in opposite

direction of previous one and lands. Takeoff must be done from a semicircular launching zone measuring 10

feet in radius.

Fig. 11: operational area in the competition

IX. FINANCIAL ASSESMENT It was one of the major aims to manufacture the UAV with minimum cost beside the unique design

requirement and the project team was successful with that desired. The cost for different parts of the UAV in the

following table:

Table 10: Financial cost in the project

ITEM COST

Motor(1200Kv) $16

ESC $10

Servo(4) $15

Battery(11.1V, 1350 mAh) $40

Connector $2

Wire, Cable $1

Remote(Transmitter & Receiver) $90

XPS(Polystyrene) Foam $6 per piece

Depron Foam $1.2 per piece

Filament tape $5

Plywood $2

Push rod $2

X. CONCLUSION The aircraft was conceptualized under a set of design requirements. However the conceptual design

converted into a real aircraft with preliminary design and detail design processes [6] with the series of limitation

as discussed. The design approach gave the right way to apply practical knowledge in real life application.

Basically the main aim was to fly the UAV with heavy payload for good scoring. Beside this aim the team tried

to keep the cost as low as possible which is the main target for the engineers now-a-days. There was a great

problem the team faced with the unavailability of the manufacturing material and electric components in local

market. These unavailable things were taken from Singapore. Through a systematic procedure finally the

aircraft was manufactured and then launched in the sky for a series of flight tests. After facing many problems through a long time (almost 2 months) the final design was manufactured for the competition. The final

launching was done successfully at the competition site (12-14 April) of SAE Aero Design Competition at

California, USA. The following figure is the successful flight in the competition. This had been a unique

experience of the project group in conceptualizing, designing, fabricating and finally participating and

qualifying in international competition with a good scoring, thereby making a history in Bangladesh and MIST

in particular.



Fig. 12: successful flight of the UAV

REFERENCES [1] John D Anderson Jr., Introduction to flight (7 West Patel Nagar, New Delhi, Tata McGraw-Hill Publishing Company Limited, Fifth

edition).

[2] L J Clancy. Aerodynamics (Melborne, Pitman Publishing Limited)

[3] Kermode, Mechanics of flight (Hong Kong, Pearson Education Limited, 2006)

[4] John D. Anderson, Jr. Fundamentals of aerodynamics (1221 Avenue of the Americas, New York, McGraw-Hill, 3rd

edition)

[5] John D. Anderson, Jr. Aircraft performance and design (7 West Patel Nagar, New Delhi, Tata McGraw-Hill Education Private

Limited, Fifth edition)

[6] Daniel P Raymer. Aircraft Design: a conceptual approach (Washington D.C., American Institute of Aeronautics and Astronautics,

Inc.)

[7] Jonathan How, Ellis King and Yoshiaki Kuwata, Flight Demonstrations of Cooperative Control for UAV Teams, AIAA 3rd

"Unmanned Unlimited" Technical Conference, Workshop and Exhibit . Chicago, Illinois, September 2004.

[8] HaiYang Chao, YongCan Cao and YangQuan Chen, Autopilots for Small Unmanned Aerial Vehicles: A Survey, International

Journal of Control, Automation, and Systems, 8(1), 2010, 36-44.

[9] Selcuk Bayraktar, Georgios E. Fainekos and George J. Pappas, Experimental Cooperative Control of Fixed-Wing Unmanned Aerial

Vehicles, 43rd IEEE Conference on Decision and Control, Atlantis, Paradise Island, Bahamas, December, 2004.

[10] Lee Nicholai's White Paper (students.sae.org/competitions/aerodesign/rules/aero_nicolai.doc)

[11] UIUC Airfoil Coordinate Database. (http://www.ae.uiuc.edu/mselig/ads/coord_database.html)

http://students.sae.org/competitions/aerodesign/rules/aero_nicolai.doc

http://www.ae.uiuc.edu/mselig/ads/coord_database.html

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